US20100071377A1 - Combustor Apparatus for Use in a Gas Turbine Engine - Google Patents

Abstract

A combustor apparatus for use in a gas turbine engine. The combustor apparatus includes a liner, a flow sleeve, and a fuel injection system. The liner includes an inner volume, wherein a portion of the inner volume defines a main combustion zone. The flow sleeve receives compressed air, is positioned radially outward from the liner, and includes a forward end and an aft end. The fuel injection system is coupled to the flow sleeve and provides fuel into the inner volume of the liner downstream from the main combustion zone. The fuel injection system includes a fuel manifold and a fuel dispensing structure. The fuel manifold is coupled to the flow sleeve and includes a cavity for receiving fuel. The fuel dispensing structure is associated with the cavity and distributes fuel from the cavity to the liner inner volume.

Description

CROSS REFERENCE TO RELATED APPLICATION
• This application is A CONTINUATION-IN-PART APPLICATION of and claims priority to U.S. patent application Ser. No. 12/233,903, (Attorney Docket No. 2008P16712US), filed on Sep. 19, 2008,” entitled “COMBUSTOR APPARATUS IN A GAS TURBINE ENGINE” the entire disclosure of which is incorporated by reference herein.
• This invention was made with U.S. Government support under Contract Number DE-FC26-05NT42644 awarded by the U.S. Department of Energy. The U.S. Government has certain rights to this invention.
FIELD OF THE INVENTION
• The present invention relates to a combustor apparatus in a gas turbine engine comprising a fuel injection system coupled to a flow sleeve for providing fuel to an inner volume of a liner.
BACKGROUND OF THE INVENTION
• In gas turbine engines, fuel is delivered from a source of fuel to a combustion section where the fuel is mixed with air and ignited to generate hot combustion products defining working gases. The working gases are directed to a turbine section. The combustion section may comprise one or more stages, each stage supplying fuel to be ignited.
SUMMARY OF THE INVENTION
• In accordance with a first embodiment of the present invention, a combustor apparatus is provided for use in a gas turbine engine. The combustor apparatus comprises a liner, a flow sleeve, and a fuel injection system. The liner comprises an inner volume, wherein a portion of the inner volume defines a main combustion zone. The flow sleeve receives compressed air, is positioned radially outward from the liner, and comprises a forward end and an aft end. The fuel injection system is coupled to the flow sleeve and provides fuel into the inner volume of the liner downstream from the main combustion zone. The fuel injection system comprises a fuel manifold and a fuel dispensing structure. The fuel manifold is coupled to the flow sleeve and includes a cavity for receiving fuel. The fuel dispensing structure is associated with the cavity and distributes fuel from the cavity to the liner inner volume.
• The fuel dispensing structure may comprise a fuel injector that distributes fuel from the fuel manifold cavity to the liner inner volume.
• The fuel injector may extend radially inwardly from the fuel manifold into an opening formed in the liner.
• The combustor apparatus may include a sliding seal member having a bore for receiving the fuel injector. The seal member may be positioned over the opening in the liner through which the fuel injector extends. The liner opening may be sized so as to be larger than an outer peripheral dimension of the fuel injector. The sliding seal member may be movably coupled to the liner so as to accommodate relative movement between the fuel injector and the liner while substantially preventing fluid leakage out from the liner opening.
• The cavity may comprise an annular channel.
• The fuel dispensing structure may include an annular array of fuel injectors that distribute fuel from the annular channel to the liner inner volume.
• The combustor apparatus may include a fuel supply structure that delivers fuel from a source of fuel to the fuel injection system. The fuel supply structure may be located radially outwardly from the flow sleeve.
• The fuel manifold may be integrally formed with the flow sleeve aft end.
• The fuel manifold may be separately formed from and affixed to the flow sleeve aft end.
• The flow sleeve may comprise a section of reduced stiffness adjacent to the fuel manifold.
• At least one gap may be formed between the fuel injection system and the liner to permit compressed air to flow through the at least one gap into the flow sleeve.
• In accordance with a second embodiment of the invention, a combustor apparatus is provided for use in a gas turbine engine. The combustor apparatus comprises a liner, a flow sleeve, and a fuel injection system. The liner comprises an inner volume, wherein a portion of the inner volume defines a main combustion zone. The flow sleeve receives compressed air, is positioned radially outward from the liner, and comprises a forward end and an aft end. The fuel injection system is associated with the flow sleeve, and provides fuel into the inner volume of the liner downstream from the main combustion zone. The fuel injection system comprises a fuel manifold and fuel dispensing structure. The fuel manifold is coupled to the flow sleeve and includes a channel that receives a fuel. The fuel dispensing structure is associated with the channel that distributes fuel from the channel to the liner inner volume. The fuel dispensing structure comprises a plurality of fuel injectors that extend radially inwardly from the fuel manifold into a plurality of openings in the liner.
• In accordance with a third embodiment of the invention, a combustor apparatus is provided for use in a gas turbine engine. The combustor apparatus comprises a liner, a flow sleeve, a first fuel injection system, a first fuel supply structure, a second fuel injection system, and a second fuel supply structure. The liner comprises an inner volume, wherein a portion of the inner volume defines a main combustion zone. The flow sleeve receives compressed air, is positioned radially outward from the liner, and comprises a forward end and an aft end. The first fuel injection system is associated with the flow sleeve, and the first fuel supply structure is in fluid communication with a source of fuel for delivering fuel from the source of fuel to the first fuel injection system. The second fuel injection system is associated with the flow sleeve aft end, and the second fuel supply structure is in fluid communication with the source of fuel for delivering fuel from the source of fuel to the second fuel injection system. The second fuel injection system provides fuel into the inner volume of the liner downstream from the main combustion zone and comprises a fuel manifold and a fuel dispensing structure. The fuel manifold is coupled to the flow sleeve aft end and includes a cavity in fluid communication with the second fuel supply structure. The fuel dispensing structure is associated with the cavity and distributes fuel from the cavity to the liner inner volume.
• The cavity may comprise a channel and the fuel dispensing structure may comprise a plurality of fuel injectors that extend radially inwardly from the fuel manifold into respective openings formed in the liner.
BRIEF DESCRIPTION OF THE DRAWINGS
• While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:
• FIG. 1 is a sectional view of a gas turbine engine including a plurality of combustors according to an embodiment of the invention;
• FIG. 2 is a side cross sectional view of one of the combustors shownFIG. 1; and
• FIG. 2A is a side cross sectional view of the pre-mix fuel injector assembly illustrated inFIG. 2 shown removed from the combustor.
• FIG. 3 is a sectional view of a gas turbine engine including a plurality of combustors having fuel supply systems according to another embodiment of the invention;
• FIG. 4 is a side cross sectional view of one of the combustors illustrated inFIG. 3 incorporating a fuel supply system according to an embodiment of the invention;
• FIG. 5 is a perspective view of the fuel supply system illustrated inFIG. 4 shown removed from the combustor;
• FIG. 6 is a perspective view of a pair of fuel supply structures of the fuel supply system illustrated inFIG. 4 shown removed from the combustor and from a combustor shell of the fuel supply system;
• FIG. 7 is a side cross sectional view of a combustor incorporating a fuel supply system according to another embodiment of the invention;
• FIG. 8 is an enlarged cross sectional view so as to illustrate a cross sectional portion in a radial and circumferential plane of a seal structure included in the combustor illustrated inFIG. 7;
• FIG. 9 is an enlarged cross sectional view so as to illustrate a cross sectional portion in a radial and circumferential plane of a fuel injector structure according to another embodiment of the invention;
• FIG. 10 is an enlarged cross sectional view so as to illustrate a cross sectional portion in a radial and circumferential plane of a fuel injector structure according to yet another embodiment of the invention; and
• FIG. 11 is an enlarged cross sectional view so as to illustrate a cross sectional portion in a radial and circumferential plane of a fuel injector structure according to yet another embodiment of the invention.
DETAILED DESCRIPTION OF THE INVENTION
• In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, specific preferred embodiments in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
• Referring toFIG. 1, agas turbine engine10 is shown. Theengine10 includes acompressor section12, acombustion section14 including a plurality of combustors13, also referred to herein as “combustion apparatuses,” and aturbine section16. Thecompressor section12 inducts and pressurizes inlet air which is directed to the combustors13 in thecombustion section14. Upon entering the combustors13, the compressed air from thecompressor section12 is pre-mixed with a fuel in a pre-mixing passage18 (seeFIG. 2). The pre-mixed fuel and air then flows into a combustion chamber14A where it is mixed with fuel from one or moremain fuel injectors15 and a pilot fuel injector17 (seeFIG. 2) and ignited to produce a high temperature combustion gas flowing in a turbulent manner and at a high velocity. The main andpilot fuel injectors15,17 are also referred to herein as “a first fuel injection system.” Thestructure11 for supplying fuel to the main andpilot fuel injectors15,17 from a fuel source is referred to herein as “a first fuel supply structure.” The combustion gas then flows through atransition26 to theturbine section16 where the combustion gas is expanded to provide rotation of aturbine rotor20 as shown inFIG. 1.
• Referring toFIG. 2, thepre-mixing passage18 is defined by a pre-mixfuel injector assembly19, also referred to herein as “a fuel injection system” or “a second fuel injection system,” comprising aflow sleeve22, also referred to herein as “a combustor shell,” surrounding aliner29 of the combustion chamber14A. Theflow sleeve22 may have a generally cylindrical configuration and may comprise anannular sleeve wall32 that defines thepre-mixing passage18 between thesleeve wall32 and theliner29. Theflow sleeve22 may be manufactured in any manner, such as, for example, by a casting procedure. Further, thesleeve wall32 may comprise a single piece or section of material or a plurality of joined individual pieces or sections, and may be formed from any material capable of operation in the high temperature and high pressure environment of thecombustion section14 of theengine10, such as, for example, stainless steel or carbon steel, and in a preferred embodiment comprises a steel alloy including chromium.
• As shown inFIG. 2, thesleeve wall32 includes a radiallyouter surface34, a radiallyinner surface35, aforward end36, and anaft end38 opposed from theforward end36. Theforward end36 is affixed to acover plate25, i.e., with bolts (not shown). Theaft end38 defines an air inlet from a combustor plenum21 (seeFIG. 1), which receives the compressed air from thecompressor section12 via a compressor section exit diffuser23 (seeFIG. 1). The radiallyouter surface34 is defined by a substantially cylindricalfirst wall section32A that extends axially between theforward end36 and theaft end38. In the embodiment shown, the radiallyinner surface35 is partially defined by thefirst wall section32A and is partially defined by asecond wall section32B. Thesecond wall section32B comprises a conical shapedportion41 and cylindrical shapedportion39. Thesecond wall section32B is affixed to and extends from thefirst wall section32A at aninterface40, as may be further seen inFIG. 2A. Thesecond wall section32B may be affixed to thefirst wall section32A by any conventional means, such as by welding.
• As seen inFIGS. 2 and 2A, theconical portion41 of thesecond wall section32B defines a transition between two inner diameters of thesleeve wall32 extending axially between theforward end36 and theaft end38. Specifically, theconical portion41 transitions between a first, larger inner diameter D₁, located adjacent to theforward end36, and a second, smaller inner diameter D₂, located adjacent to the aft end38 (seeFIG. 2A). It is understood that thesleeve wall32 may have a substantially constant diameter if desired, or the diameter D₂of theaft end38 could be greater than the diameter D₁of theforward end36.
• Referring toFIGS. 2 and 2A, acavity42 is defined in thesleeve wall32 adjacent to the sleeve wall aftend38 between the first andsecond wall sections32A,32B. In the preferred embodiment, thecavity42 comprises a first portion defining atransition chamber44 and a second portion defining an annularfuel supply chamber46, but may comprise any number of portions, including a single portion.
• In the illustrated embodiment, thefuel supply chamber46 is separated from thetransition chamber44 by aweb member48 extending radially between the first andsecond wall sections32A,32B and dividing thecavity42 into thetransition chamber44 and thefuel supply chamber46. It should be noted that although theweb member48 is illustrated as comprising a separate piece of material attached to the first andsecond wall sections32A,32B, theweb member48 could also be provided as integral with either or both of the first andsecond wall sections32A,32B of thesleeve wall32.
• The annularfuel supply chamber46 comprises anannular channel46A formed in thesleeve wall32 and defines a fuel flow passageway for supplying fuel around the circumference of thesleeve wall32 for distribution to thepre-mixing passage18. Theannular channel46A may be formed in thesleeve wall32 by any suitable method, such as, for example, by bending or forming the end of thesleeve wall32 or by machining theannular channel46A into thesleeve wall32. In the embodiment shown, theannular channel46A preferably extends circumferentially around theentire sleeve wall32, but may extend around only a selected portion of thesleeve wall32. Optionally, thefuel supply chamber46 may be provided with a thermallyresistant sleeve58 therein, i.e., a sleeve formed of a material having a high thermal resistance. Additional description of theannular channel46A and the thermallyresistant sleeve58 may be found in U.S. patent application Ser. No. 12/180,637, (Attorney Docket No. 2005P15727US), filed on Jul. 28, 2008 entitled “INTEGRAL FLOW SLEEVE AND FUEL INJECTOR ASSEMBLY,” the entire disclosure of which is incorporated by reference herein.
• Referring toFIG. 2, theflow sleeve22 further comprises afuel feed passageway24 provided for receiving afuel supply tube49, whichtube49 is also referred to herein as “a fuel supply structure” or “a second fuel supply structure” and also defines a “fuel supply element,” that is in fluid communication with a source offuel50 and extends through anaperture25A in thecover plate25. As may be further seen inFIG. 2A, thefuel feed passageway24 is defined by aU-shaped cover structure27 that is affixed to theinner surface35 of thesleeve wall32, such as by welding, for example, and is further defined by a slot or opening47 (FIG. 2) defined in thesecond wall section32B at theconical portion41. Thecover structure27 isolates thefuel supply tube49 from the hot gases flowing through thepre-mixing passage18 by substantially preventing the hot gases from entering thefuel feed passageway24. Hence, thefuel supply tube49 provides fluid communication for conveying fuel between the source offuel50 and thefuel supply chamber46 of thecavity42 by passing through theaperture25A in thecover plate25, through thefuel feed passageway24, including theopening47, and through thetransition chamber44 of thecavity42. TheU-shaped cover structure27 and the first andsecond wall sections32A,32B defining thetransition chamber44 are also referred to herein as “shield structure.”
• Referring toFIG. 2A, thefuel supply tube49 is affixed to theweb member48, for example, by welding, such that afluid outlet24A of thefuel supply tube49 is in fluid communication with thefuel supply chamber46 of thecavity42 via an aperture48A formed in theweb member48. Preferably, as most clearly shown inFIG. 2A, thefuel supply tube49 may include a series ofbends49A,49B or circumferential direction shifts within thetransition chamber44 of thecavity42, so as to provide thefuel supply tube49 with an S-shape. As shown inFIG. 2A, the S-shaped fuel supply tube has a first section extending along a first path having a component in an axial direction, a second section extending along a second path having a component in a circumferential direction, and a third section extending along a third path having a component in the axial direction. Thebends49A,49B may reduce stress to thefuel supply tube49 caused by a thermal expansion and contraction of thefuel supply tube49 and theflow sleeve22 during operation of theengine10, accommodating relative movement between thefuel supply tube49 and thesleeve wall32, such as may result from thermally induced movement of one or both of thefuel supply tube49 andsleeve wall32. Thefuel supply tube49 may be secured to thesleeve wall32 at various locations withfasteners52A,52B, illustrated herein by straps, as seen inFIGS. 2 and 2A. It should be understood that other types of fasteners, allowing any combination of free and constrained degrees of freedom could be used and could be employed in different locations than those illustrated inFIGS. 2 and 2A.
• Referring toFIGS. 2 and 2A, afuel dispensing structure54 is associated with theannular channel46A and, in the preferred embodiment, comprises anannular segment46B of thesleeve wall32 adjacent theaft end38. In the embodiment shown, theannular segment46B is provided as a separate element affixed in sealing engagement over theannular channel46A to form a radially inner boundary for theannular channel46A, and is configured to distribute fuel into thepre-mixing passage18. For example, theannular segment46B may be welded to thesleeve wall32 at first and second welds (not shown) on opposed sides of theannular channel46A at an interface between theannular segment46B and thesleeve wall32 to create a substantially fluid tight seal with thesleeve wall32. It should be noted that other means may be provided for affixing theannular segment46B to thesleeve wall32 and that theannular segment46B of thefuel dispensing structure54 could be formed integrally with thesleeve wall32. Thefuel dispensing structure54 is further described in the above-noted U.S. patent application Ser. No. 12/180,637 (Attorney Docket No. 2005P15727US).
• Thefuel dispensing structure54 further includes a plurality offuel distribution apertures56 formed in theannular segment46B. In a preferred embodiment, thefuel distribution apertures56 comprise an annular array of openings or through holes extending through theannular segment46B. Thefuel distribution apertures56 may be substantially equally spaced in the circumferential direction, or may be configured in other patterns as desired, such as, for example, a random pattern. Thefuel distribution apertures56 are adapted to deliver fuel from thefuel supply chamber46 to thepre-mixing passage18 at predetermined circumferential locations about theflow sleeve22 during operation of theengine10. The number, size and locations of thefuel distribution apertures56, as well as the dimensions of thefuel supply chamber46, are preferably configured to deliver a predetermined flow of fuel to thepre-mixing passage18 for pre-mixing the fuel with incoming air as the air flows to the combustion chamber14A.
• Since thecover structure27 is formed integrally with theflow sleeve22, the possibility of damage to thefuel supply tube49, which may occur during manufacturing, maintenance, or operation of theengine10, for example, may be reduced by the present design. Further, thecover structure27 and thetransition chamber44 of thecavity42 prevent direct contact and provide a barrier for thefuel supply tube49 from vibrations that would otherwise be imposed on thefuel supply tube49 by the gases flowing through the pre-mixing passage28. Accordingly, damage caused to thefuel supply tube49 by such vibrations is believed to be avoided by the current design.
• Moreover, theaft end38 of thesleeve wall32 provides a relatively restricted flow area at the entrance to thepre-mixing passage18 and expands outwardly in the flow direction producing a venturi effect, i.e., a pressure drop, inducing a higher air velocity in the area of thefuel dispensing structure54. The higher air velocity in the area of thefuel dispensing structure54 facilitates heat transfer away from theliner29 and substantially prevents flame pockets from forming between thesleeve wall32 and theliner29, which could result in flames attaching to and burning holes in thesleeve wall32, theliner29, and/or any other components in the vicinity. Further, while the pressure drop provided at theaft end38 of thesleeve wall32 is sufficient to obtain the desired air velocity increase adjacent to thefuel dispensing structure54, a substantial pressure is maintained along the length of theflow sleeve22 in order to limit the production of NO_xin the fuel/air mixture between thesleeve wall32 and theliner29.
• Theweb member48 located at theaft end38 of thesleeve wall32 forms an I-beam structure with the first andsecond wall sections32A,32B to strengthen and substantially increase the natural frequency of theflow sleeve22 away from the operating frequency of the combustor13. For example, the operating frequency of the combustor13 may be approximately 300 Hz, and the natural frequency of theflow sleeve22 is increased by the I-beam stiffening structure to approximately 450 HZ. Hence, damaging resonant frequencies in theflow sleeve22 are substantially avoided by the increase in the natural frequency provided by the present construction.
• A portion of a can-annular combustion system114, constructed in accordance with a further embodiment of the present invention, is illustrated inFIG. 3. Thecombustion system114 forms part of agas turbine engine110. Thegas turbine engine110 further comprises acompressor112 and aturbine118. Air enters thecompressor112, where it is compressed to an elevated pressure and delivered to thecombustion system114, where the compressed air is mixed with fuel and burned to create hot combustion products defining a working gas. The working gases are routed from thecombustion system114 to theturbine118. The working gases expand in theturbine118 and cause blades coupled to a shaft and disc assembly to rotate.
• The can-annular combustion system114 comprises a plurality ofcombustor apparatuses116 and a like number ofcorresponding transition ducts120. Thecombustor apparatuses116 andtransition ducts120 are spaced circumferentially apart so as to be positioned within and around an outer shell orcasing110A of thegas turbine engine10. Eachtransition duct120 receives combustion products from itscorresponding combustor apparatus116 and defines a path for those combustion products to flow from thecombustor apparatus116 to theturbine118.
• Only asingle combustor apparatus116 is illustrated inFIG. 4. Each of thecombustor apparatuses116 forming part of the can-annular combustion system114 may be constructed in the same manner as thecombustor apparatus116 illustrated inFIG. 4. Hence, only thecombustor apparatus116 illustrated inFIG. 4 will be discussed in detail here.
• Thecombustor apparatus116 comprises a combustor shell126 (also referred to herein as a flow sleeve) coupled to theouter casing110A of thegas turbine engine110 via acover plate135, seeFIG. 4. Thecombustor apparatus116 further comprises aliner128 coupled to thecover plate135 viasupports128A, a firstfuel injection system116A, firstfuel supply structure116A₁, a secondfuel injection system116B and secondfuel supply structure116B₁. Thecombustor shell126 may comprise anannular shell wall130. Anair flow passage124 is defined between theshell wall130 and theliner128 and extends up to thecover plate135.
• As shown inFIG. 4, theshell wall130 includes a radiallyouter surface131, a radiallyinner surface132, aforward end133, and anaft end134 opposite theforward end133. Theforward end133 is affixed to thecover plate135 of theengine110, i.e., with bolts (not shown). Thecover plate135 is coupled to theouter casing110A viabolts136A, seeFIG. 4. Theaft end134 defines a first inlet into theair flow passage124. Compressed air generated by thecompressor112 passes through anexit diffuser138 andcombustor plenum137 prior to passing through theaft end134 into theair flow passage124, seeFIG. 3.
• In the illustrated embodiment, theshell wall130 comprises a plurality ofapertures139 defining a second inlet into theair flow passage124. Further compressed air generated by thecompressor112 passes from outside theshell wall130 into theair flow passage124 via theapertures139. It is understood that the percentage of air that passes into theair flow passage124 through theapertures139 versus that which passes through the first inlet defined by theaft end134 of theshell wall130 can be configured as desired. For example, 100% of the air may pass into theair flow passage124 at the first inlet defined by theaft end134, in which case theapertures139 would not be necessary. Or, nearly all of the air may pass into theair flow passage124 through theapertures139, although it is understood that other configurations could exist. Theapertures139 are designed, for example, to condition and/or regulate the flow around the circumference of theshell wall130 such that if it is found that more/less air is needed at a certain circumferential location, then theapertures139 at that location could be enlarged/reduced in size andapertures139 in other locations could be reduced/enlarged in size accordingly. It is contemplated that theapertures139 may be arranged in rows or in a random pattern and, further, may be located elsewhere in theshell wall130. Further, theshell wall130 may include a radially inwardly taperedportion140 adjacent to theaft end134 thereof, as shown inFIGS. 4 and 5.
• The firstfuel injection system116A comprises apilot nozzle200 attached to thecover plate135 and a plurality ofmain fuel nozzles202 also attached to thecover plate135, seeFIG. 4. The firstfuel supply structure116A₁comprising firstfuel inlet tubes216 coupled to thepilot nozzle200 and themain fuel nozzles202 as well as to a fuel source152. Thefuel inlet tubes216 receive fuel from the fuel source152 and provide the fuel to the pilot andmain fuel nozzles200 and202. The fuel from the pilot andmain fuel nozzles200 and202 is mixed with compressed air flowing through theair flow passage124 and ignited in a combustion chamber ormain combustion zone114A within theliner128 creating combustion products defining a working gas.
• The secondfuel injection system116B is located downstream from the firstfuel injection system116A and comprises anannular manifold170 coupled to the shell wall aftend134, such as by welding, seeFIGS. 4-6. A plurality offuel injectors172 extend radially inwardly from themanifold170. Thefuel injectors172 extend into an inner volume of theliner128 so as to inject fuel, viaopenings172A, into theliner128 at a location downstream from themain combustion zone114A, seeFIG. 4. It is noted that injecting fuel in two fuel injection locations, i.e., via the firstfuel injection system116A and the secondfuel injection system116B, may reduce the production of NOx by thecombustion system114. For example, since a significant portion of the fuel, e.g., about 15-25% of the total fuel supplied by the first and secondfuel injection systems116A,116B, is injected in a location downstream of thecombustion chamber114A, i.e., by the secondfuel injection system116B, the amount of time that the combustion products are at a high temperature is reduced as compared to combustion products resulting from the ignition of fuel injected by the firstfuel injection system116A. Since NOx production is increased by the elapsed time the combustion products are at a high combustion temperature, combusting a portion of the fuel downstream of thecombustion chamber114A reduces the time the combustion products resulting from the fuel provided by the secondfuel injection system116B are at a high temperature such that the amount of NOx produced by thecombustion system114 may be reduced. Thefuel injectors172 may be substantially equally spaced in the circumferential direction about the manifold170, or may be configured in other patterns as desired, such as, for example, a random pattern. The number, size and locations of thefuel injectors172 andopenings172A, as well as the dimensions of theannular manifold170, may vary.
• The secondfuel supply structure116B₁communicates with theannular manifold170 of the secondfuel injection system116B and the fuel source152 so as to provide fuel from the fuel source152 to the secondfuel injection system116B, seeFIG. 4. The secondfuel supply structure116B₁comprises first and secondfuel supply elements144A,144B, asecond inlet tube316 and athird inlet tube318, seeFIGS. 4-6. The firstfuel supply element144A comprises a firsttubular line156 having first, second andthird sections156A,156B and156C. Thefirst section156A is coupled to thecover plate135 and communicates with a fitting314A, which, in turn, communicates with thesecond inlet tube316. Thesecond inlet tube316 is coupled to the fuel source152. Thefirst section156A of the firsttubular line156 extends away from thecover plate135 along a first path P₁having a component in an axial direction, which axial direction is indicated by arrow A inFIG. 5. Thesecond section156B extends along a second path P₂, which second path P₂has a component in a circumferential direction. The circumferential direction is indicated by arrow C inFIG. 5. In the illustrated embodiment, the second path P₂extends about 90 degrees to the first path P₁and through an arc of about 180 degrees. It is contemplated that the second path P₂may extend through any arc within the range of from about 15 degrees to about 180 degrees. Thethird section156C extends along a third path P₃having a component in the axial direction A. In the illustrated embodiment, the third path P₃extends about 90 degrees to the second path P₂and is generally parallel to the first path P₁. Thethird section156C is coupled to aninlet170A of themanifold170. Hence, fuel flows from the fuel source152, through thesecond inlet tube316, the fitting314A, the firstfuel supply element144A and into themanifold inlet170A so as to provide fuel to themanifold170.
• The secondfuel supply element144B comprises a secondtubular line158 having fourth, fifth andsixth sections158A,158B and158C. Thefourth section158A is coupled to thecover plate135 and communicates with a fitting (not shown), which, in turn, communicates with thethird inlet tube318. Thethird inlet tube318 is coupled to the fuel source152. Thefourth section158A of the secondtubular line158 extends away from thecover plate135 along a fourth path P₄having a component in the axial direction A. Thefifth section158B extends along a fifth path P₅, which fifth path P₅has a component in the circumferential direction C. In the illustrated embodiment, the fifth path P₅extends about 90 degrees to the fourth path P₄and through an arc of about 180 degrees. It is contemplated that the fifth path P₅may extend through any arc within the range of from about 15 degrees to about 180 degrees. Thesixth section158C extends along a sixth path P₆having a component in the axial direction A. In the illustrated embodiment, the sixth path P₆extends about 90 degrees to the fifth path P₅and is generally parallel to the fourth path P₄. Thesixth section158C is coupled to aninlet170B of themanifold170. Hence, fuel flows from the fuel source152, through thethird inlet tube318, the fitting, the secondfuel supply element144B and into themanifold inlet170B so as to provide further fuel to themanifold170.
• As shown inFIGS. 2-4, the third andsixth sections156C and158C of the first and secondtubular lines156 and158 includeangled parts156D and158D. Theangled parts156D and158D causeend parts156E and158E of the third andsixth sections156C and158C to bend inwardly so as to follow the radially inwardly taperedportion140 of theshell wall130.
• During operation of thecombustor apparatus116, thecombustor shell wall130 may thermally expand and contract differently, i.e., a different amount, from that of theannular manifold170, which is coupled to theaft end134 of thecombustor shell wall130, as well as differently from that of the secondfuel supply structure116B₁. This is because the fuel flowing through the secondfuel supply structure116B₁and theannular manifold170 functions to cool the secondfuel supply structure116B₁and theannular manifold170. Hence, during operation of thecombustor apparatus116, thecombustor shell wall130 may reach a much higher temperature than theannular manifold170 and the secondfuel supply structure116B₁. Further, thecombustor shell wall130 may be made from a material with a coefficient of thermal expansion different from that of the material from which theannular manifold170 and/or the secondfuel supply structure116B₁are made. The different coefficients of thermal expansion and different operating temperatures may result in different rates and amounts of thermal expansion and contraction during combustor apparatus operation and, hence, may contribute to differing amounts of thermal expansion and contraction between thecombustor shell wall130 and theannular manifold170 and/or the secondfuel supply structure116B₁. Because the first and secondtubular lines156 and158 defining the firstfuel supply elements144A and1448 have angled configurations, i.e., the second andfifth sections156B and158B extend substantially laterally to the first,third sections156A,156C and the fourth,sixth sections158A,158C, the first and secondtubular lines156 and158 are capable of deflecting as thecombustor shell wall130 and theannular manifold170/secondfuel supply structure116B₁thermally expand and contract differently. Hence, internal stresses within the first and secondtubular lines156 and158, which may normally occur ifsuch lines156 and158 had only a linear configuration, do not occur or occur at a limited amount during operation of thecombustor apparatus116.
• In the illustrated embodiment, ashield structure141 is affixed to the radiallyouter surface131 of theshell wall130, seeFIGS. 4 and 5. Theshield structure141 may be formed separately from and affixed to theshell wall130, such as by welding, for example, or may be formed integrally with theshell wall130. Further, theshield structure141 may comprise one or more separate elements that are coupled together to form theshield structure141. In the embodiment shown, theshield structure141 comprises an annular member having a generally U-shaped cross section that extends completely around theshell wall130. However, it is understood that theshield structure141 may extend around only a selected portion or portions of theshell wall130 and may have any suitable shape.
• Theshield structure141 defines a protective casing having aninner cavity142, seeFIG. 4. In the illustrated embodiment, theshield structure141 includes first andsecond inlet apertures146A and146B and first andsecond outlet apertures148A and148B. The firsttubular line156 passes through the first inlet andoutlet apertures146A and148A such that the second section1568 of the firsttubular line156 is located within theinner cavity142 of the shield structure. The secondtubular line158 passes through the second inlet andoutlet apertures146B and148B such that thefifth section158B of the secondtubular line158 is also located within theinner cavity142 of the shield structure. The second andfifth sections156B and158B of the first and secondtubular lines156 and158 extend generally transverse to the axial direction at which high velocity compressed air from the compressor passes along and near theouter surface131 of thecombustor shell wall130 and through theair flow passage124. Theshield structure141 functions to shield or protect the second andfifth sections156B and158B of the first and secondtubular lines156 and158 from impact by the high velocity compressed air moving along and near theouter surface131 of thecombustor shell wall130 and passing through theair flow passage124. If left exposed to the high velocity compressed air, the high velocity air could apply undesirable forces to the second andfifth sections156B and158B of the first and secondtubular lines156 and158, which forces may damage the first andsecond lines156 and158 or create undesirable vibrations in thelines156 and158.
• The first and secondtubular lines156 and158 may be secured to theshell wall130 or theshield structure141. In the illustrated embodiment, the second andfifth sections156B and158B of the first and secondtubular lines156 and158 are secured to theshield structure141 at various locations withfasteners166, seeFIGS. 4 and 5. Thefasteners166 preferably restrain the first and secondtubular lines156 and158 from vibration while allowing a limited amount of motion in the fore-to-aft direction to permit thermal expansion/contraction of the first and secondtubular lines156 and158, which, as noted above, may occur differently from that of theshell wall130.
• Acombustor apparatus1216 constructed in accordance with yet a further embodiment of the present invention is illustrated inFIG. 7. Each of a plurality combustor apparatuses forming part of a can-annular combustion system may be constructed in the same manner as thecombustor apparatus1216 illustrated inFIG. 7.
• Thecombustor apparatus1216 comprises a combustor shell226 (also referred to herein as a flow sleeve) coupled to anouter casing210A of agas turbine engine210 via acover plate235, seeFIG. 7. Thecombustor apparatus1216 further comprises aliner228 coupled to thecover plate235 viasupports228A, a firstfuel injection system216A, firstfuel supply structure216A₁, a secondfuel injection system216B and secondfuel supply structure216B₁. Thecombustor shell226 may comprise anannular shell wall230. Anair flow passage224 is defined between theshell wall230 and theliner228 and extends up to thecover plate235.
• As shown inFIG. 7, theshell wall230 includes a radiallyouter surface231, a radiallyinner surface232, aforward end233, and anaft end234 opposite theforward end233. Theforward end233 is affixed to thecover plate235 of theengine210, i.e., with bolts (not shown). Thecover plate235 is coupled to theouter casing210A viabolts236A, seeFIG. 7. Theaft end234 defines a first inlet into theair flow passage224. Compressed air generated by a compressor passes through an exit diffuser and combustor plenum prior to passing through theaft end234 into theair flow passage224.
• Theshell wall230 may include a radially inwardly taperedportion240, which, in the illustrated embodiment, includes theaft end234, seeFIG. 7. As will be discussed further below, in the illustrated embodiment, the taperedportion240 is less stiff than an adjacentmain portion1230 of theshell wall230. The reduction in stiffness of the taperedportion240 may result by forming the taperedportion240 with a thickness less than a thickness of themain portion1230 or by forming the taperedportion240 from a material which is less resistant to deformation than a material used to form themain portion1230. The reduction in stiffness of the taperedportion240 may also result from the formation of a plurality ofapertures239 in the taperedportion240, which apertures239 define a second inlet for the compressed air to enter into theair flow passage224. Hence, further compressed air generated by the compressor passes from outside theshell wall230 into theair flow passage224 via theapertures239.
• It is understood that the percentage of air that passes into theair flow passage224 through theapertures239 versus that which passes through the first inlet defined by theaft end234 of theshell wall230 can be configured as desired. For example, 100% of the air may pass into theair flow passage224 at the first inlet defined by theaft end234, in which case theapertures239 would not be necessary. Or, nearly all of the air may pass into theair flow passage224 through theapertures239, although it is understood that other configurations could exist. Theapertures239 are designed, for example, to condition and/or regulate the flow around the circumference of theshell wall230 such that if it is found that more/less air is needed at a certain circumferential location, then theapertures239 at that location could be enlarged/reduced in size andapertures239 in other locations could be reduced/enlarged in size accordingly. It is contemplated that theapertures239 may be arranged in rows or in a random pattern and, further, may be located elsewhere in theshell wall230.
• The firstfuel injection system216A comprises apilot nozzle300 attached to thecover plate235 and a plurality ofmain fuel nozzles302 also attached to thecover plate235, seeFIG. 7. The firstfuel supply structure216A₁comprises firstfuel inlet tubes317 coupled to thepilot nozzle300 and themain fuel nozzles302 as well as to afuel source252. Thefuel inlet tubes317 receive fuel from thefuel source252 and provide the fuel to the pilot andmain fuel nozzles300 and302. The fuel from the pilot andmain fuel nozzles300 and302 is mixed with compressed air flowing through theair flow passage224 and ignited in a combustion chamber ormain combustion zone214A within theliner228 creating combustion products defining hot working gases.
• The secondfuel injection system216B is located downstream from the firstfuel injection system216A and comprises a manifold270 coupled to the shell wall aftend234, such as by welding. It is also contemplated that the manifold270 may be formed as an integral part of theshell wall230. Hence, the manifold270 is structurally independent of theliner228, whichliner228, as will be discussed further below, typically operates at a much higher temperature than theshell wall230 and themanifold270. Hence, thermally induced stresses, which might result if the manifold270 is coupled directly to theliner228, are substantially reduced or eliminated.
• The manifold270 comprises aninner cavity271 for receiving fuel. In the illustrated embodiment, the manifold270 is annular; hence, theinner cavity271 in the manifold270 defines an annular channel. A plurality offuel injectors272 extend radially inwardly from the manifold270 and define a fuel dispensing structure. In theFIG. 8 embodiment, the manifold270 comprises outer and inner radially spaced apartwalls270A and270B. Eachfuel injector272 passes throughbores1270A and1270B in thewalls270A and270B and may be welded or otherwise held in position to one or both of thewalls270A and270B. Eachfuel injector272 comprises circumferential and radial bores272A, which communicate with the manifoldinner cavity270A so as to define a path for fuel to pass from the manifoldinner cavity270A into, through and out from thefuel injector272. Eachfuel injector272 extends through a corresponding one of a plurality ofopenings1228, seeFIG. 8, formed in theliner228 so as to inject fuel into an inner volume of theliner228 at a location downstream from themain combustion zone214A, seeFIG. 7. The fuel dispensing structure may be defined by one or a plurality of thefuel injectors272.
• As noted above, theaft end234 defines a first inlet into theair flow passage224. It is also noted that a plurality ofgaps1229, seeFIG. 8, extend radially between the manifold270 and theliner228, wherein eachgap1229 extends generally circumferentially betweenadjacent fuel injectors272. As shown by the dashed lines inFIG. 8, radial dimensions of thegaps1229 may be adjusted by changing the configuration of theinner wall270B of themanifold270. By changing the radial dimensions of thegaps1229, the amount of compressed air permitted to flow through the first inlet into theair flow passage224 can be controlled, i.e., increased or decreased, as a function of the size of thegaps1229.
• In one alternative embodiment illustrated inFIG. 9, eachfuel injector2272 passes through abore3270B in aninner wall2270B of a manifold2272 and may be welded in position to thatinner wall2270B. Further, an area of theinner wall2270B near thebore3270B is shaped so as to enlargegaps2229 between theliner228 and theinner wall2270B of themanifold2272. In a further alternative embodiment illustrated inFIG. 10, eachfuel injector3272 is threaded into a threadedbore4270B in aninner wall3273B of themanifold3270.
• In the illustrated embodiment, eachliner opening1228 is larger in size than an outer peripheral dimension of itscorresponding injector272. For example, if theinjector272 is generally cylindrical in shape with a generally circular cross section having a diameter D₁, then a diameter D₂of itscorresponding liner opening1228 is larger than the injector diameter D₁, seeFIG. 8.
• During operation of thecombustor apparatus1216, the manifold270 andfuel injectors272 may be cooled by fuel passing through them, depending upon the temperature of the fuel, but are heated by compressed air passing over them, which compressed air is provided by the compressor. During start-up and operation of thecombustor apparatus1216, the manifold270 andfuel injectors272 may heat up to a temperature within the range of from about 400° F. to about 800° F., theshell wall230 may heat up to a temperature within the range of from about 400° F. to about 800° F., and theliner228 may heat up to a temperature in excess of 1600° F. Consequently, the temperature of the manifold270 andfuel injectors272 may be slightly less than or approximately equal to the temperature of theshell wall230, such that severe thermal gradients or thermal changes between the manifold270/fuel injectors272 and theshell wall230 may not occur. However, during combustor apparatus operation, the temperatures of the manifold270, thefuel injectors272 and theshell wall230 are much lower than the temperature of theliner228, through which hot working gases pass. Consequently, theliner228 may shift relative to theinjectors272 and vice versa during start up, operation and shut-down of thecombustor apparatus1216. Because theliner openings1228 are oversized relative to theinjectors272, some amount of movement of theliner228 relative to theinjectors272 and vice versa, which movement occurs due to changing temperatures, may be accommodated such that theinjectors272 and theliner228 do not contact one another.
• As noted above, the taperedportion240 is less stiff than the adjacentmain portion1230 of theshell wall230. Thus, the taperedportion240 may accommodate differences in thermal expansion, such as in the radial direction, between the manifold270 and theshell wall230, which differences in thermal expansion may be caused by the manifold270 being at a slightly lower temperature than theshell wall230, e.g., up to about 300° F. less. For example, during operation of thecombustor apparatus1216, it is believed that themain portion1230 of theshell wall230 may expand radially a greater amount than the manifold270, i.e., the shell wall main portion diameter may expand a greater amount than the diameter of themanifold270. It is believed that the taperedportion240 will flex or otherwise accommodate these thermally induced differences in the diameters of themain portion1230 and the manifold270 so as to minimize thermal-induced stresses between theshell wall230 and themanifold270. The lower temperature of the manifold270 relative to theshell wall230 may be attributed to the fuel flowing through the manifold270, which fuel may have a temperature in a range from about 70° F. to about 800° F. It is also believed that theliner228 may expand radially a greater amount than the manifold270, i.e., the liner diameter may expand a greater amount than the diameter of themanifold270. As a result, the radial dimensions of thegaps1229 between theliner228 and the manifold270 will decrease, causing thefuel injectors272 to extend further through corresponding seal member bores402 (discussed further below) and thecorresponding liner openings1228. Thus, in an embodiment, the seal members bores402 and thefuel injectors272 are configured such that relative radial movement, i.e., radial sliding, can occur therebetween. The lower temperature of the manifold270 relative to the liner may be attributed to the fuel flowing through the manifold270 and the hot working gases flowing through theliner228, which working gases may have a temperature of up to about 2800° F.
• So as to minimize the amount of working gases escaping through theliner openings1228, a plate-like slidingseal member400 is associated with eachliner opening1228, seeFIG. 8. The slidingseal member400 comprises abore402 for receiving acorresponding fuel injector272. The size of thebore402 is only slightly larger than the diameter D₁of theinjector272 such that little or no hot working gases pass between theinjector272 and theseal member400. However, the bore size must be large enough to accommodate radial movement of itscorresponding injector272, as noted above. Theseal member400 extends over itscorresponding liner opening1228 so as to cover theopening1228. Theseal member400 is movably or slidably coupled to theliner228 so as to allow it to move with itsfuel injector272 relative to theliner228. As noted above, theliner228 may move relative to thefuel injectors272 and vice versa as the temperatures of theshell wall230, theliner228, the manifold270 and thefuel injectors272 vary relative to one another during operation of thecombustor apparatus1216. In the illustrated embodiment, clips404, e.g., fourclips404, are fixed to theliner228, which define with theliner228oversized recesses406 for receiving edges of theseal member400, e.g., four edges of a generally square orrectangular seal member400. Therecesses406 capture theseal member400 so as to couple it to theliner228, yet allow theseal member400 to move relative to theliner228 and itscorresponding liner opening1228, seeFIG. 8. In an alternative embodiment, a plate-like slidingseal member4000 is associated with eachliner opening4228, seeFIG. 11. In this embodiment, the slidingseal member4000 comprises abore4020 for receiving acorresponding fuel injector4272. The size of thebore4020 is only slightly larger than a diameter of theinjector4272 such that little or no hot working gases pass between theinjector4272 and theseal member4000. However, the bore size must be large enough to accommodate radial movement of itscorresponding injector4272, as noted above. Theseal member4000 is movably or slidably coupled to theliner228 so as to allow it to move with itsfuel injector4272 relative to theliner228. Specifically, in the embodiment shown inFIG. 11, acircumferential tooth4040 defines theliner opening4228 and extends toward theseal member4000. Theliner tooth4040 is received in a slot defined by radially inner and radiallyouter teeth4050A and4050B of theseal member4000. As shown inFIG. 11, theliner opening4228 is oversized, such that theseal member4000 can slide axially and/or circumferentially with respect to theliner228, while staying engaged with thetooth4040. That is, theseal member teeth4050A,4050B capture thetooth4040 so as to couple theseal member4000 to theliner228, yet allow theseal member4000 to move relative to theliner228 and itscorresponding liner opening4228, seeFIG. 11.
• It is noted that injecting fuel at two axially spaced apart fuel injection locations, i.e., via the firstfuel injection system216A and the secondfuel injection system216B, may reduce the production of NOx by thecombustor apparatus1216. For example, since a significant portion of the fuel, e.g., about 15-30% of the total fuel supplied by the firstfuel injection system216A and the secondfuel injection system216B, is injected at a location downstream of themain combustion zone214A, i.e., by the secondfuel injection system216B, the amount of time that the second combustion products are at a high temperature is reduced as compared to first combustion products resulting from the ignition of fuel injected by the firstfuel injection system216A. Since NOx production is increased by the elapsed time the combustion products are at a high combustion temperature, combusting a portion of the fuel downstream of themain combustion zone214A reduces the time the combustion products resulting from the second portion of fuel provided by the secondfuel injection system216B are at a high temperature, such that the amount of NOx produced by thecombustor apparatus1216 may be reduced.
• Thefuel injectors272 may be substantially equally spaced in the circumferential direction, or may be configured in other patterns as desired, such as, for example, a random pattern. Further, the number, size, and location of thefuel injectors272 andcorresponding liner openings1228 may vary depending on the particular configuration of thecombustor apparatus1216 and the amount of fuel to be injected by the secondfuel injection system216B.
• The secondfuel supply structure216B₁communicates with themanifold270 of the secondfuel injection system216B and thefuel source252 so as to provide fuel from thefuel source252 to the secondfuel injection system216B, seeFIG. 7. The secondfuel supply structure216B₁may comprise the same elements and be constructed in the same manner as the secondfuel supply structure116B₁illustrated inFIG. 4-6. It is noted that the secondfuel supply structure216B₁is located adjacent theouter surface231 of theshell wall230 and, hence, is protected from the high velocity compressed air passing into and through theair flow passage224, which comprises the majority of the compressed air coming from the compressor to thecombustor apparatus1216.
• While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.

Claims (20)

1. A combustor apparatus for use in a gas turbine engine comprising:

a liner comprising an inner volume, wherein a portion of said inner volume defines a main combustion zone;

a flow sleeve for receiving compressed air, said flow sleeve positioned radially outward from said liner and comprising a forward end and an aft end; and

a fuel injection system coupled to said flow sleeve, said fuel injection system providing fuel into said inner volume of said liner downstream from said main combustion zone, said fuel injection system comprising;

a fuel manifold coupled to said flow sleeve and including a cavity for receiving fuel; and

a fuel dispensing structure associated with said cavity, said fuel dispensing structure distributing fuel from said cavity to said liner inner volume.

2. The combustor apparatus according toclaim 1, wherein said fuel dispensing structure comprises a fuel injector that distributes fuel from said fuel manifold cavity to said liner inner volume.

3. The combustor apparatus according toclaim 2, wherein said fuel injector extends radially inwardly from said fuel manifold into an opening formed in said liner.

4. The combustor apparatus according toclaim 3, further comprising a sliding seal member having a bore for receiving said fuel injector, said seal member being positioned over said opening in said liner through which said fuel injector extends, said liner opening being sized so as to be larger than an outer peripheral dimension of said fuel injector, said sliding seal member being movably coupled to said liner so as to accommodate relative movement between said fuel injector and said liner while substantially preventing fluid leakage out from said liner opening.

5. The combustor apparatus according toclaim 1, wherein said cavity comprises an annular channel.

6. The combustor apparatus according toclaim 5, wherein said fuel dispensing structure includes an annular array of fuel injectors that distribute fuel from said annular channel to said liner inner volume.

7. The combustor apparatus according toclaim 1, further comprising a fuel supply structure that delivers fuel from a source of fuel to said fuel injection system, said fuel supply structure located radially outwardly from said flow sleeve.

8. The combustor apparatus according toclaim 1, wherein said fuel manifold is integrally formed with said flow sleeve aft end.

9. The combustor apparatus according toclaim 1, wherein said fuel manifold is separately formed from and affixed to said flow sleeve aft end.

10. The combustor apparatus according toclaim 1, wherein said flow sleeve comprises a section of reduced stiffness adjacent to said fuel manifold.

11. The combustor apparatus according toclaim 1, wherein at least one gap is formed between said fuel injection system and said liner to permit compressed air to flow through said at least one gap into said flow sleeve.

12. A combustor apparatus for use in a gas turbine engine comprising:

a liner comprising an inner volume, wherein a portion of said inner volume defines a main combustion zone;

a flow sleeve for receiving compressed air, said flow sleeve positioned radially outward from said liner and comprising a forward end and an aft end; and

a fuel injection system associated with said flow sleeve, said fuel injection system providing fuel into said inner volume of said liner downstream from said main combustion zone, said fuel injection system comprising;

a fuel manifold coupled to said flow sleeve and including a channel receiving a fuel; and

fuel dispensing structure associated with said channel that distributes fuel from said channel to said liner inner volume, said fuel dispensing structure comprising a plurality of fuel injectors that extend radially inwardly from said fuel manifold into a plurality of openings in said liner.

13. The combustor apparatus according toclaim 12, further comprising a plurality of sliding seal members, at least one of said sliding seal members having a bore for receiving a corresponding one of said fuel injectors, said one seal member being positioned over a corresponding one of said openings in said liner and being movably coupled to said liner so as to move with said one fuel injector relative to said liner.

14. The combustor apparatus according toclaim 12, further comprising a fuel supply structure that delivers fuel from a source of fuel to said fuel injection system, said fuel supply structure located radially outwardly from said flow sleeve.

15. The combustor apparatus according toclaim 12, wherein said fuel manifold is integrally formed with said flow sleeve aft end.

16. The combustor apparatus according toclaim 12, wherein said fuel manifold is separately formed from and affixed to said flow sleeve aft end.

17. A combustor apparatus for use in a gas turbine engine comprising:

a liner comprising an inner volume, wherein a portion of said inner volume defines a main combustion zone;

a flow sleeve for receiving compressed air, said flow sleeve positioned radially outward from said liner and comprising a forward end and an aft end;

a first fuel injection system associated with said flow sleeve;

a first fuel supply structure in fluid communication with a source of fuel for delivering fuel from said source of fuel to said first fuel injection system;

a second fuel injection system associated with said flow sleeve aft end;

a second fuel supply structure in fluid communication with said source of fuel for delivering fuel from said source of fuel to said second fuel injection system;

said second fuel injection system providing fuel into said inner volume of said liner downstream from said main combustion zone, said second fuel injection system comprising;

a fuel manifold coupled to said flow sleeve aft end and including a cavity in fluid communication with said second fuel supply structure; and

a fuel dispensing structure associated with said cavity, said fuel dispensing structure for distributing fuel from said cavity to said liner inner volume.

18. The combustor apparatus according toclaim 17, wherein:

said cavity comprises a channel; and

said fuel dispensing structure comprises a plurality of fuel injectors that extend radially inwardly from said fuel manifold into respective openings formed in said liner.

19. The combustor apparatus according toclaim 18, further comprising a plurality of sliding seal members, at least one of said sliding seal members having a bore for receiving a corresponding one of said fuel injectors, said one seal member being positioned over a corresponding one of said openings in said liner and being movably coupled to said liner so as to move with said one fuel injector relative to said liner.

20. The combustor apparatus according toclaim 17, wherein said flow sleeve comprises a section of reduced stiffness adjacent to said fuel manifold.

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